Modern gas turbine engines, and more specifically turbofans for use in aviation, provide power by compressing air using a compressor, adding fuel to this compressed air, combusting this mixture such that it expands through the blades of a turbine and exhausting the produced gases.
With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In modern engines, the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air cooling of these airfoil components. During its passage through the engine, the mean temperature of the gas stream decreases as power is extracted. Therefore, the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.
FIG. 2 shows an isometric view of a typical single stage cooled turbine. Cooling air flows are indicated by arrows.
Internal convection and external films are the prime methods of cooling the gas path Components—airfoils, platforms, shrouds and shroud segments etc. High-pressure turbine nozzle guide vanes 31 (NGVs) consume the greatest amount of cooling air on high temperature engines. High-pressure blades 32 typically use about half of the NGV flow. The intermediate-pressure and low-pressure stages downstream of the HP turbine use progressively less cooling air.
The high-pressure turbine airfoils are cooled by using high pressure air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while gas temperatures can be in excess of 2100 K.
The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air effectively.
Ever increasing gas temperature levels combined with a drive towards flatter combustion radial profiles, in the interests of reduced combustor emissions, have resulted in an increase in local gas temperature experienced by the extremities of the blades and vanes, and the working gas annulus endwalls.
This has led to the concept of a shroudless turbine, where the shroud is eliminated (or substantially reduced), allowing a better management of cooling flow in response to a flatter temperature profile, and a reduction in loading on the disc due to a reduction of parasitic mass.
In an enshrouded turbine, the turbine consists of a disc, rotating about the central shaft of the engine, and a plurality of blades extending radially out from the disc towards the engine casing of the engine. Expansion through the turbine causes its blades to rotate at high speed. The blades of the turbine rotate closely to, and within, an annular assembly of seal segments that form a seal segment ring around the tips of the turbine blades.
As the disc and the blades of the turbine rotate, they are subject to considerable centrifugal force and temperatures. The centrifugal force and high temperature cause the turbine to extend in the radial direction and this can cause “rubbing” as the tips of the blades come into contact with the seal segments.
As the turbine rotates, the distance between the tips of the blades and the seal segments is known as the tip clearance. It is desirable for the tips of the turbine blades to rotate as close to the seal segments without rubbing as possible because as the tip clearance increases, the efficiency of the turbine decreases, as a portion of the expanded gas flow will pass through the tip clearance. This is known as over-tip leakage.
Steps have been taken in order to minimise the over tip leakage. Currently some engines are fitted with an abradable coating on the inside of the seal segment, against which the blades can rub. This means that the blades will cut a groove in this lining, to form a seal against the casing.
Conventionally, the seal segment is cooled by impingement. Coolant is contained within an annular duct formed by the inner wall of the engine casing and the outer surface of the ring seal segments, which impinges the coolant onto the back face of the seal segment inner wall. This method requires a large flow of coolant to sufficiently cool the abradable layer and keep it at an acceptable temperature due to the thermal resistance of the seal segment wall between the coolant and the radially inner surface of the abradable material.
Another approach to cooling is effusion cooling. Coolant is discharged from an outboard annular duct through one or more very small diameter radial ducts, which extend through the seal segment and abradable coating. The coolant can therefore cool the seal segment and abradable coating, being in direct contact with both of these components, and finally be impinged on the turbine blade tips.
GB2009329A describes a turbine wheel shroud of porous abradable material, incorporating transverse, non-porous, divisions to prevent axial flow losses of the cooling air along the length of the shroud.